1. Field of the Invention
The present invention generally relates to supersonic vehicles, and more particularly, to a supersonic ground vehicle having substantially reduced aerodynamic pressure drag.
2. Prior Art
When an object travels through the atmosphere at speeds greater than the speed of sound shock waves are formed. The generation of these shock waves requires a significant amount of power and results in considerable aerodynamic wave pressure drag on the object.
The classical approach to designing aircraft having reasonably low supersonic drag includes the incorporation of design attributes such as a pointed nose having a proper included angle, swept wings, and "area ruled" volume geometry. Such designs are utilized on all modern supersonic aircraft.
Currently, the only supersonic ground vehicles in use are rocket test sleds. These sleds are used to evaluate either aircraft crew ejection equipment or kinetic energy weapons. The aerodynamic design of existing rocket sleds do not utilize any special means for reducing drag at supersonic speeds apart from the classical approaches used in the design of supersonic aircraft. Existing rocket sleds overcome the large drag associated with their designs by using high powered rocket engines to propel them to supersonic speeds.
Existing rocket sleds have disadvantages associated with their high-powered rocket engines. For instance, rocket engines must utilize highly explosive materials to create the thrust needed to obtain supersonic speeds in a high drag environment. Using such highly explosive materials obviously creates dangerous explosive hazards. In addition, using highly explosive materials creates high operating costs due to the costs associated with the amount of material necessary to obtain the required thrust. In addition, flight at these speeds create noise levels not suitable for the human environment. In view of these disadvantages, supersonic ground transportation has been considered economically infeasible because the power required to propel such vehicles is prohibitive and because of noise control.
A method using favorable aerodynamic interference has been proposed to reduce or eliminate drag in supersonic aircraft. This method requires the implementation of a two-dimensional or axisymmetric "Supersonic Biplane" configuration which causes the shock waves generated by each wing to favorably interact with each other. Such configurations also have been suggested for use in the area of interference between wings and engine nacelles, where similar wave interactions could theoretically be effected with positive results. However, these low drag supersonic aircraft configurations have never been realized or used on aircraft due to their inherent structural complexity and because such vehicles must necessarily operate over a range of angles-of-attack and speeds which adds considerable complexity to the design of the favorably interfering elements.
The "Supersonic Biplane" concept was first developed by A. Busemann in the early 1900's, and thus, is known as the "Busemann Biplane" effect. Busemann's studies were centered around two-dimensional wedges and were presented in the absence of all frictional forces (viscosity). In the real world, the presence of frictional forces prevent the optimum condition of zero drag. However, because viscous effects are confined to very small regions close to solid surfaces, the presence of frictional forces do not prevent great reductions in and true minimizations of aerodynamic wave drag. The explanation that follows is based on the two-dimensional, frictionless theory of Busemann, however, Busemann's theory can be extended to explain why applicant's invention reduces aerodynamic drag in three-dimensional cross-sectional configurations in real world viscous environments.
During supersonic flight, a vehicle travels at speeds greater than the speed of sound. At such speeds, air is compressed just ahead of the vehicle. This compression region, or shock wave, is actually a high pressure sound wave that creates what is known as a "sonic boom". The creation of this disturbance is commonly called "breaking the sound barrier". Since this disturbance travels as a wave it is susceptible to the physical laws of wave cancellation and reinforcement. These physical laws are the basic mechanisms behind the "Busemann Biplane" effect.
To simplify the description of the "Busemann Biplane" effect let us suppose the air is moving at supersonic speed and the vehicle is stationary. In this condition, compare the air flow past a Busemann type two-dimensional upper/lower wedge configuration (see FIG. 1), and the air flow past a classically designed supersonic aircraft (see FIG. 2). The arrows in FIGS. 1 and 2 indicate the direction of the air flow which is traveling at a velocity (V) which is greater that the speed of sound (V.sub.sound).
In the case of an aircraft in flight (FIG. 2), shock waves (sound waves) are created by the sudden change in direction of the incoming flow path caused by the nose of the aircraft. The wedges which form the nose are at an angle which causes the incoming flow to change direction, slow down and become compressed. The angle that the shock wave makes with the horizontal incoming flow is dependent upon the wedge angle and the velocity of the incoming flow. If the speed, wedge angle, or distance between the wedges is changed a variety of wave patterns are created.
To minimize the drag of the wedge geometry one needs to eliminate or minimize the exiting shock waves. The minimization of these waves implies that the outgoing air flow is returned back to the same flow condition that existed in the upstream undisturbed incoming flow field. One such flow pattern is shown in FIG. 1, where the wedge shaped configuration creates a shock wave which forms a symmetric "X" pattern in the forward and aft portions of the configuration.
In order to understand why this pattern is created the flow at the apex of the wedge must be examined. In FIG. 1, the incoming air flow intersects the wedges at the forward edges. The slope of the forward edges turn the flow and create shock waves. Since the angle of the forward edges with respect to the air flow is less than 90.degree. the flow slows and the air is compressed. The generated shock waves then impinge at the apex of the wedges and are reflected away from the wedges. If the geometry is selected (as shown) such that the wedges are symmetric forward and aft, then the reflected waves will again intersect the wedges at the trailing edges. The intersection with the trailing edges causes the flow to be turned afterward. This time, however, the angle is greater than 90.degree., thus, the air expands and the flow speeds up, thereby eliminating the shock wave. This is exactly the opposite effect that occurs to the incoming flow.
If all conditions are met, the incoming and the outgoing flow are at the same condition and the air flow does not undergo any lasting change, hence, aerodynamic drag is effectively eliminated. However, in the real world this effect does not occur perfectly and these small changes in air characteristics cause minor increases in drag.